The Southern California Selene Group withdrew from competition in June 2008.
At the dawn of the Space Age, conventional wisdom had it that geostationary communication satellites would be too complex to be commercially practical. Harold Rosen, who formed and led the team that designed and built the first successful geostationary satellite, the spin-stabilized Syncom launched in 1963, put that misconception to rest.
Then as now, Dr. Rosen has felt that many space exploration missions could likewise be simplified. The announcement of the Google Lunar XPRIZE has created the opportunity to demonstrate this.
To this end, and to compete for the grand prize, he has pulled together a volunteer group of talented, experienced and enthusiastic team members to form the Southern California Selene Group. These team members constitute some of the finest engineering minds in the space business, with over 130 space-related patents to their credit and active participation in over 500 space missions.
The architecture for the “Spirit of Southern California” spacecraft will combine the elegantly simple control and communication systems used in some of the earliest communications satellites with the latest in electronic and sensor technology.
This approach taken by the Southern California Selene Group can be succinctly summarized as “an elegantly simple design that can be implemented quickly and inexpensively.”
Dr. Harold Rosen, Team Leader. Harold has earned worldwide recognition for his pioneering work in the field of communications satellites and is widely recognized as “the father of the geostationary satellite” in that he formed and led the team that designed and built the first successful geostationary satellite, Syncom, and subsequently, as Vice President, went on to help build the world’s largest communications satellite business at Hughes Aircraft Company. Dr. Rosen has received the 1995 National Academy of Engineering’s Draper Prize, the 1990 Arthur C. Clarke Award (presented by the President of Sri Lanka), the 1985 National Medal of Technology (presented by President Reagan), the 1985 Communications and Computing Prize from NEC, the 1982 Alexander Graham Bell Medal and the 1976 Ericsson International Prize in Communications (presented by the King of Sweden). In 2003, he was inducted into the National Inventors Hall of Fame. In addition to the above, he has received numerous other awards and honors, among them the 1992 Design News Special Achievement Award, the 2003 Discover magazine Innovation Award, and the ISCe 2004 Lifetime Achievement Award. A holder of over seventy-five space related patents, he is a Fellow of the IEEE and the AIAA. He is a Distinguished Alumnus of Caltech, from which he received his PhD in electrical engineering and aeronautics. Dr. Rosen now consults for Boeing in the design of new satellite systems.
Deborah Castleman, Associate Team Leader. Deborah, formerly retired, was a satellite systems engineer at the Space and Communications Group of Hughes Aircraft Company, a space and defense policy analyst at RAND, the Deputy Assistant Secretary of Defense for Command, Control and Communications in the first Clinton Administration, and Vice President of Rosen Motors, which was developing a hybrid electric powertrain for automotive use. Deborah holds an MS in Electrical Engineering from the California Institute of Technology, an MA in International Studies from the Claremont Graduate School (now Claremont Graduate University) and a BS in Electrical and Electronic Engineering from the California State Polytechnic University at Pomona. Earlier, she served in the U.S. Air Force as an avionics technician for the F-111 aircraft.
Ron Symmes, Project Manager. Ron is President of Symmetrics, which provides management consulting services to numerous aerospace and entrepreneurial start-up companies. He is also Executive Vice President and co-founder of Global Aerospace, a technical service company to the satellite industry, and founding member and an owner of Legacy Engineering, a supplier of senior level contract labor to the aerospace industry. Ron retired from Hughes Space and Communications Group as Vice President of Operations, where he was a satellite system engineer, a program manager for several international satellite programs, and Division Manager for system engineering and operations. In 1997 he received the AIAA Space Systems Award for “outstanding leadership in the design, development, and orbit operations of the HS-601 communication satellite product line.” Ron received his BSEE and MSEE from Northeastern University in Boston.
Rex Ridenoure, Deputy Project Manager. Rex is CEO and co-founder of Ecliptic Enterprises Corporation. He has nearly 30 years of experience in space-mission architecting, mission engineering and operations, spacecraft systems and space project management. For the first 20 years of his career he worked on more than a dozen pioneering space projects: multiple Shuttle-launched geostationary communications satellite missions at Hughes Space & Communications Group, the Hubble Space Telescope at Lockheed Missiles and Space Company, microsatellites at Utah State University, and deep-space missions such as Voyager and Deep Space 1 at JPL, where he was also deeply involved with multiple lunar and low-cost deep-space mission studies and proposals. In the late 1990s, he successfully transitioned into the entrepreneurial space arena, holding leadership roles at Microcosm, SpaceDev and BlastOff! Corporation. For individual and team contributions, Rex has been a recipient or co-recipient of several major professional space awards, most notably (with three others) a 1999 Aviation Week & Space Technology Laurel Award for playing a key role in the 1998 salvage of the stranded HGS-1 comsat using a novel orbit method that made HGS-1 the first — and still only — commercial spacecraft to reach the Moon's distance. He holds an MS from Caltech in Aeronautics and BS from Iowa State University in Aerospace Engineering.
Dorian Challoner, Lead Systems Engineer. Dorian works at Boeing and is an expert with three decades of experience in the design and development of space robotics and spacecraft guidance, navigation and control systems. He has made significant contributions to the control system architecture, design and analysis of the Communications Technology Satellite, Shuttle Remote Manipulator and Boeing satellites since 1979. He holds thirty patents related to satellite attitude dynamics and control, and to inertial micro-electromechanical systems (MEMS). Among these are key patents related to the Boeing disc resonator gyroscope. Dorian has also conducted research in the field of fluid dynamics for spinning satellites and inertial MEMS. He currently is the chief scientist for the leading DARPA program in that field. Dorian received a BASc in Engineering Science and a MASc in Controls Systems Engineering from the University of Toronto.
Robert Rosen, Systems Engineer. Robert, who works at Raytheon part time and consults for Pentadyne, is an expert in the design of advanced digital radars, and is especially proficient in multi-body orbital analysis and control system design and simulation. Earlier, at the Space and Communications Group of Hughes Aircraft Company, one of the many critical analytic contributions Robert made was to determine the maximum allowable time-to-transition between passive spin stability of the HS-376 satellite and the onset of active nutation control as the communications payload was despun. This important result was subsequently used, during this critical phase in orbital attitude dynamics, for the ground-breaking and hugely successful (over 60 put in orbit!) HS-376 series of satellites. Robert holds an MS in Physics from UC Berkeley, and a BS in Physics from UCLA.
Dr. Alois Wittmann, Systems Engineer. Before he retired in 2003, Al was Chief Technologist and Senior Technical Fellow of the Boeing Company (formerly Hughes Space and Communications Group). During his over 37 years with the company, he was engaged in the design, testing, and flight performance evaluation of complete spacecraft, as well as spacecraft component research and development programs. He developed the mechanical design architectures for the Hughes extended power spinners HS-376, the HS-393 and Intelsat VI wide-body satellite. For the later wide-body spacecraft, he invented the Frisbee type shuttle cargo deployment, and introduced the integral orbit acquisition stage and carrier cradle structure for optimal shuttle utilization. When the satellite production changed to the body stabilized designs of HS-601 and HS-702, he contributed to their mechanical and structure system optimization, as well as the application of the Xenon Ion electrical propulsion for an industry leading technology. Dr. Wittmann engineered the first flexible substrate roll-up solar array and conducted its experimental space flight evaluation. Application solar array designs followed under his design lead such as Kevlar and graphite unfolding panel arrays. He promoted the use of enhanced composite material technology, developing low temperature curing epoxy systems for extremely stable antenna and microwave devices. Dr. Wittmann received his education in Germany and his doctorate engineering degree at the Technical University of Munich in 1961. He holds 15 patents, and has received many awards, among them the 1981 Hughes Aircraft Lawrence A. Hyland Patent Award (Hughes' highest invention award), the 1992 Hughes Telecommunication and Space Sector patent award, the 1999 Hughes Chairman’s Honor award, and the 2002 Boeing Technical Excellence Award.
Dr. John Smay, Systems Engineer. John is the author of 11 U. S. Patents in spacecraft attitude control technology, and has been a contributing participant in more than 100 space missions. In 1999, he retired as Chief Technologist, reporting to the Company President, at the Hughes Space and Communications Group (now Boeing Space Systems). Prior to this, John was the Assistant Laboratory Manager of the Guidance and Control Systems Laboratory, where he contributed to the design and analysis of most SCG spacecraft and space payload programs, including Intelsat IVA, Comstar, HS-333, HS-318, HS-350, HS-111, MMB I-III, HS-601 and HS-702. He was the lead attitude control design engineer for the GMS, the HS-376 and HS-393 product lines (more than 55 spacecraft over 18 years), and actively participated in more than 80 commercial and military spacecraft launch missions, including the first spacecraft retrieval (Westar/Palapa in 1984) and several major anomaly events such as the SBS F4 imbalance, the HS-393 platform release failure, a flat spin recovery and the Intelsat VI F-1 capture. While at Hughes, John was awarded the Space and Communications Group Invention Award in 1988, the Space Sector Patent Award in 1992, the L.A. Hyland Patent Award in 1996 (Hughes' highest invention award), and the Telecommunications and Space Company Technical Excellence Award in 1996. Dr. Smay is also the author and presenter of 13 technical papers on automatic control and spacecraft attitude control subjects. Due to his international reputation, he was asked by the President of China Great Wall Industry Corp. to serve on a Long March failure review team. More recently, he has consulted for Hughes, Boeing, SES Astra (Luxembourg), Bristol Aerospace (Canada) and XM Radio. He holds a BS, MS, and PhD in Electrical Engineering from the University of Colorado.
Brian Bliss, Systems Engineer. Brian is a satellite payload development engineer at Boeing, assigned to the Global Positioning System (GPS) program. His focus to date has been high-reliability embedded logic design and real-time software for radiation environments. His first integrated circuit design was launched into orbit in October 2007 aboard the U.S. Air Force’s first Wideband Global Satcom satellite. He previously worked as an embedded software developer for the startup Starbak Communications, Inc. He holds dual BS degrees from Purdue University, one in Computer Engineering and the other in Aeronautical and Astronautical Engineering, where he specialized in dynamics and control. He is currently pursuing an MS in Systems Engineering and an MBA at Loyola Marymount University in Los Angeles.
Patty Pun, Systems Engineer. Patty is a Design Engineer at Qualcomm, and her current job responsibility is software development, focusing on instrument control and data acquisition, to support measurements of electromagnetic emission profiles of ASICs and Printed Circuit Boards. She is proficient in Labview, Matlab and other software development tools. Previously, she was a research programmer on a robot sensor network, with a focus on camera and robot control, motion detection, real time image streaming; and a research programmer for 3D modeling of Volumetric Visualization of the Convection-generated Stresses in Earth. She earned a M.S. in Electrical and Computer Engineering (ECE), and B.S. in ECE with Robotics Minor from Carnegie Mellon University.
Daniel Geng, Systems Engineer. Dan is a digital payload engineer at Boeing, and is proficient in digital hardware design, digital image processing, video/ image compression and error correction coding. He serves as a Responsible Engineering Authority (REA) on the Global Positioning System (GPS) program. He helped design and analyze the digital payload architecture of numerous commercial, civil, and military satellites. He received an Engineering Excellence Award from Boeing for his work on GPS Block-IIF, which is scheduled for launch starting August 2008. He earned a B.S. degree from California Institute of Technology in Electrical Engineering, and currently pursuing a M.S. in Computer Science at University of Southern California.
Stan Kent, Systems Engineer. As the founder of the Viking Fund, Stan is one of the pioneers of privately funded space exploration in that he raised the necessary funds to keep the Viking lander on Mars operating when budget cuts threatened the ongoing mission. Stan is a rocket propulsion expert who works at Boeing. He previously worked for Aerojet, Lockheed, Hughes and NASA. At Hughes he was head of the Propulsion Subsystem Engineering department and oversaw the propulsion development of the HS-601, HS-601HP and HS-376HP. He conducted specialized thruster testing to support the Intelsat VI re-boost mission, where Space Shuttle astronauts salvaged a satellite left in a lower orbit due to a booster malfunction. Stan is a veteran of some thirty communication satellite missions and holds a patent on an optimized method for analyzing transfer orbit performance. He holds a BS and MS in Astronautics from Stanford University. and was awarded the Herman Oberth gold medal by the International Astronautical Federation for outstanding and original research while still a student. A prolific writer, Stan has penned nine full length novels and edited the volume “Remember the Future: the Apollo Legacy."
Philip Donatelli, Systems Engineer. Currently, Phil provides consulting services on spacecraft propulsion system design and on-orbit operation to Boeing Space Systems and other spacecraft manufacturers. He retired from Hughes Space and Communications Group after 35 years of service in 1997, where was the Manager of the Propulsion Systems Laboratory. Prior to joining Hughes, he worked at Reaction Motors Division, Thiokol Chemical Corporation designing injectors and thrust chambers for liquid propellant engines. Phil joined Hughes in the early phases of the Surveyor Lunar Lander Program as Senior Project Engineer for the Vernier Propulsion System and actively participated in all 7 Surveyor missions. Subsequently, he was the primary individual for the design and qualification of a Hughes-manufactured, long life hydrazine thruster that has flown on a number of Hughes commercial spacecraft such as the Intelsat IVA, SBS and HS-376. When the preferred choice for spacecraft propulsion systems transitioned from monopropellant to bipropellant systems, he was actively involved in the development and qualification of the propulsion system and the 22N thruster for the Intelsat VI spacecraft and a low cost 10N thruster assembly that is used for some of the HS-601 body-stabilized spacecraft configurations. He holds a BS in Mechanical Engineering from Lehigh University, Bethlehem, PA. He has 4 patents for monopropellant hydrazine thrusters and spacecraft components. In addition, he was a member of a team that was awarded the Boeing Patent and Technical Excellence Award in 2000.
Josh Rosen and Max Johnson, Graphics/ Animation. Both Josh and Max design computer-generated 3D presentations and develop software. Max also specializes in web development, while Josh specializes in creating 3D rendering software. Both currently attend high school.
SCSG System and Mission Summary, May 2008
Our system incorporates many design features of the 1960s spin-stabilized Syncom communication satellite, adding the elements needed for the transfer orbit to the moon, landing and roaming (via hopping) on the lunar surface, and implementing the electronic elements in modern, mostly digital components.
The previous soft landers on the moon, the 1960s U.S.S.R. Luna series and the U.S. Surveyor series, employed three axis body-stabilized configurations. These require more complex attitude and orbit control systems, thermal control systems, instrumentation systems, and landing systems than the optimally designed spin-stabilized system we are using. The use of modern electronic elements not available during all these earlier missions, when combined with simplifications associated with the spinning configuration, result in a total mission cost, including launch costs, of about one percent of the Surveyor costs in today’s dollars. We have filed a provisional patent covering the salient features of our design.
Our initial (February 2008) design began its mission with a launch by a SpaceX Falcon 1e of a 700 kilogram payload into a 200 kilometer altitude Earth orbit. The payload consisted of a perigee rocket stage (an ATK Thiokol Star 30), a multi-function interstage, a retro rocket (an ATK Thiokol Star 17), and the lander, which also performed the hopping function. Both the interstage and the lander were outfitted with a monopropellant hydrazine propulsion system.
The choice of a launch vehicle necessarily drives our design. SpaceX recently issued it new Falcon Lunar Capability Guide, which upgraded the performance of the Falcon 1e to provide up to 1000 kilograms of mass into low Earth orbit. This has enabled us to incorporate some design improvements (we now plan to have a 900 kilogram payload) that had not been feasible under the previous 700 kilogram payload allowance.
For this reason, we have made two major changes to our original configuration. The first is that we no longer have a liquid propulsion system on the interstage; the functions that had been performed by this system will be performed by the lander instead. This results in a greatly simplified interstage. The second is that we are using a bipropellant propulsion system instead of a monopropellant system. These improvements have allowed for more mass allocation for the electronics, and we have increased range with the potential of hopping up to 5000 meters.
Upon separation from the launch vehicle, the payload will spin itself up, from the low spin rate imparted by the launch vehicle, to 100 RPM, using small solid spin-up rockets on the interstage. The spin-stabilized perigee rocket will impart a velocity increment of 3140 meters per second to the assembly, injecting it into a 90 hour lunar transfer orbit. The spacecraft will then head for its targeted Moon landing, shortly after local lunar sunrise, near 70 degrees West, 0 degrees North in lunar coordinates.
After burnout of the perigee stage, the remaining mass will be 290 kilograms. The perigee case, along with the interstage structure, will be jettisoned. The spin axis will then be precessed to ecliptic normal attitude by pulsing the axial thrusters on the lander in a pre-programmed maneuver. This attitude, in which the sun illuminates the spinning cylinder but not its end closures, provides a benign thermal environment for all the electronic and propulsion components of the lander during its cruise to the moon. It is the most favorable attitude for the spinning sun and moon sensors used for attitude determination. It also ensures that the earth remains in the telemetry and command antenna beams for the first phase of the transfer orbit. This allows range and range rate data to be obtained as soon as the spacecraft appears in view (because the ecliptic normal attitude places the antenna in a position that is not obscured by the spacecraft) of the first telemetry and command earth station, so that the actual orbit can be determined, and the first in a series of orbit corrections can be executed. The remaining assembly now consists of the lander and the embedded retro rocket, and has a combined mass of 230 kilograms. The lander has a mass of approximately 150 kilograms, which includes 100 kilograms of bipropellant fuel, of which approximately 20 kilograms is expected to be used for orbit corrections.
These orbit corrections will occur as early as possible in order to minimize fuel consumption. They will be imparted by the axial and radial thrusters on the lander without changing the spin axis attitude. As the mission progresses, several such corrections of decreasing magnitude may be required until the desired orbit accuracy is obtained.
The spin-stabilized lander has a spinning section that contains the solar cell power supply; the propulsion system; sun and moon sensors for attitude determination; a landing radar; telemetry and communication transmitters; a command receiver; antennas; and a control processor. It also contains a retro rocket. A despun section contains the landing gear and a microwave transparent mast, which surrounds the antennas and supports the camera assembly at the top.
Shortly before the lunar encounter, the spin axis will be reoriented to the braking attitude by pulsing the lander’s axial thrusters. This reorientation maneuver (as was the case in the earlier reorientation maneuver to the ecliptic normal) is controlled by miniature rate MEMS gyros in the lander. The accuracy of these gyros is enhanced by the spin, which neutralizes their bias errors. The resulting spin axis will be within a few milliradians of the desired braking attitude. The landing site chosen results in a nearly vertical approach angle, which makes the landing simpler than other approach angles since only a small change in spin axis attitude is required during the descent. This choice of landing site also permits the use of the transfer orbit communications antenna system for the lunar operations.
After lots of intensive design work exploring the use of an all-liquid descent system, we have converged on a design whereby the retro-rocket function is instead a hybrid descent system, in which a solid rocket provides roughly half the velocity increment required for the landing phase, while the other half is provided by the lander's axial thrusters.
At an altitude of several hundred kilometers, with the approach speed approximately 2500 meters per second, the solid rocket is ignited, burns for 18 seconds, and is jettisoned. Then the descent phase using the axial thrusters begins.
The radar in the lander measures its altitude and vector velocity relative to the moon, starting at an altitude of approximately 50 kilometers. The radar beam is aimed 20 degrees away from the spin axis so that the ensuing modulation of the Doppler frequency at the spin rate measures the transverse velocity vector component. This avoids the need for multiple radars to measure all of the approach velocity components.
As we approach the moon, the horizontal velocity errors will be driven to zero by pulsing the radial thrusters. The axial thrusters, through an appropriate descent velocity versus altitude profile, control the lander to a soft landing. At about a meter altitude, the thrusters are turned off. The ensuing free fall is cushioned by the flexible landing legs and stabilized by the gyroscopic torques resulting from the spin angular momentum. The lander, except for the legs and camera system, remains spinning during the lunar phase of the mission.
Since the lander has the power supply, communication system, instrumentation, and propulsion systems needed for the roaming part of the mission, it is both mass and cost effective to roam (hop) the lander itself instead of providing for a separate roamer that would need to duplicate the functions of these systems. The cost of roaming in our approach is merely the additional fuel needed for the hops. We expect to have at least 15 kilograms of fuel available for hopping after a nominal landing.
Hopping rather than crawling enables us to travel across the surface easily, because the presence of obstacles in the path can be avoided by a hopping roamer that can leap over them in a single bound.
After the initial Mooncasts are transmitted to earth, the lander begins its hopping. A hop is accomplished by using the axial thrusters to gain altitude while the radial thrusters accelerate the craft laterally, with the spin axis mostly in a vertical attitude. This attitude may initially be slightly off vertical due to the slope of the landing site; if so, it will be corrected to vertical during the ascent. The descending portion of the hop is similar to the final (vernier) portion of the initial landing.
All communications with our craft -- telemetry, command, and wideband data from the moon -- will be via the communication system on the lander.
The telemetry and wideband data links, which share the same downlink antenna and ten watt power amplifier, will operate in the allocated space-to-earth portion of S-band. The command link will use the uplink antenna and will operate in the earth-to-space portion of S-band. The command receiver and telemetry transmitter will provide a phase coherent two-way Doppler signal for orbit determination during the 90 hour transfer orbit. We have started the process of requesting the experimental license to use these frequencies from the Federal Communications Commission.
The power transmitted by our downlink antenna permits a substantially higher data rate capability than the amount needed for the Mooncasts, and provides the ability to transmit a data stream continuously throughout the different phases of the mission. The downlink antenna provides more than sufficient gain (about 5 dB) without requiring any moving parts or electronic beam steering. We are able to do this because our orbit control system provides the velocity corrections at a constant spin axis attitude, thus the direction to the earth stays constant relative to the spin axis throughout the transfer orbit. It also is constant throughout the lunar operations, and differs from the transfer orbit case by only 20 degrees. This permits us to use an antenna with a relatively small beamwidth in any plane containing the spin axis, enhancing its gain.
This downlink antenna will be a coaxial cable fed collinear array of sleeved dipoles located inside an RF-transparent mast. Its beam pattern will be a figure of revolution around the spin axis, but it will have directivity in any plane containing the spin axis, with a beamwidth of 30 degrees. The beam center will be squinted north by 10 degrees, so that the off beam center gain loss will be the same for the transfer orbit (in which the earth is broadside to the spin axis) and lunar operations (which for the chosen 70 degrees West longitude site and a locally vertical spin axis attitude is 20 degrees north of broadside).
In contrast, the uplink antenna does not require as much gain. It will consist of a single coaxial cable fed sleeved dipole located just above the transmit antenna. In this simple arrangement, the outer conductor of the uplink antenna’s coaxial feed will be the inner conductor of the downlink antenna’s coaxial feed. An antenna similar to this one was used on the first geostationary satellite, Syncom, and on the first commercial communication satellite, Early Bird. By such use, we intend to demonstrate that some of these earlier technologies are appropriate for use in low-cost access to space today.
We plan to use the Universal Space Network to meet our telemetry and command requirements. Three longitudinally separated earth stations will be used to provide a continuous link to the spacecraft during the transit. If necessary, we will provide the equipment that has the requisite features needed for our orbit and attitude determination.
To meet the all the Mooncast requirements, and at the same time keep our system simple and elegant, we have designed our camera system from commercially available components. The camera system is located at the highest point of our lander, at the top of the mast. This enables us to look downward to the top of the lander and nearby lunar surface, as well as outward to the distant lunar surface.
The various specified Mooncast modes will be configured electronically. Our detector will be a single focal plane array whose associated electronics provides the versatility for the required self-portrait detailed images, near real time video, and high definition video. The camera itself will stare in a generally upward orientation at a tiltable mirror that provides the required elevation viewing range. The pan requirement will be met by rotating the camera/mirror assembly around the mast axis. In order for the camera to see the required minimum of 40 percent of the lander surface area, the cylindrical surface -- which contains solar cells, radial thrusters, and the sun sensor -- must be viewable. An additional mirror will be used to meet this requirement. The appropriate mechanisms for the pan, tilt and focus adjustments have been designed.
The camera is located on a non-rotating mast while the transmitters, receivers, decoders and processors are located on the spinning portion of the lander. Because of this, the power and control signals to the camera and its mirror, and the data output from the camera, have to cross a rotating interface. A slip-ring brush assembly is thus required to conduct signals back and forth. The uplink coaxial antenna feed is used for the additional function of bringing the signals to the slip rings. (Slip rings and brushes have long been used on most dual-spin communication satellites to serve a similar purpose.)
To support the required Mooncasts, the Allen Telescope Array is ideal for our needs. The ATA has 1000 square meters of capture area (from the presently installed 42 antennas) and a low front end noise temperature, which -- when combined with the noise of the moon -- gives a system noise temperature of 180K (46K from the ATA front end and 134K from the moon). In conjunction with our downlink antenna gain and transmitter power, this permits a data rate of at least 2 Megabits per second to be transmitted.
- Harold Rosen